S[)&@ O[f)@~&uO@[m~ C@~~u[~~

FINAL REPORT. VOl.UME IV

SOC SYSTEM ANAL VSIS R E (BOOK 2 OF 2)

0180-26495-4

DRl. T-1591 LINE ITEM 4 ORO MA-69i'T

(.~~ L""tt .~ q r .. \1 "'-...t w l.

111111111111111111111111111111111111111111111 NF02556

THE lI~t:I'E'NC COMPANY

HAMILTC)N STANDARD 0 DVISKX10f UNITED TECHNOLOGIES.

NASA-CR-167558 19820012330

https://ntrs.nasa.gov/search.jsp?R=19820012330 2018-06-24T20:55:59+00:00Z

0180-26495-4

SPACE OPERATIONS CENTER

SYSTEM ANALYSIS

Conducted for the NASA Johnson Space Center

Under Contra:ct NAS9-16151

FINAL REPORT

VOLUME IV

(Book 2 of 2)

SYSTEM ANALYSIS REPORT

D 180--26495-4

July ,1981 ,. /

1 , j' /

~~ /1' /sj-' ~. / / (", Ii ;f fl'. " Td~u 'Il'f1!l;((4Z;~ Approved by ''''\

BOEING AEROSPACE COMPANY

P.O. BOX 3999

Seattle, Washington 98124

Gordon R. Woodcock, SOC Study Manager

0180-26495-4

FOREWORD

The Space Operations Center System Analysis study (Contract NAS9-16151) was

initiated in June of 1980 and completed in May of 1981. This was the equivalent of

a NASA Phase A study. A separately funded Technology Assessment and

A.dvancement Plan study was conducted in parallel with the System A.nalysis Study.

These studies were managed by the Lyndon B. Johnson Space Center. The

Contracti.ng Officers Representative and Study Technical Manager is Sam Nassiff.

This study was conducted by The Boeing Aerospace Company, Large Space Systems

Group with the Hamilton Standard Division of United Technologies as subcontrac-

tor. The Boeing study manager is Gordon R. Woodcock. The Hamilton Standard

study manager is Harlan Brose.

This final report includes 8 documents:

D180-26495-1 Vol. I

D 180-26495-2 Vol. II

D180-26495-3 Vol. III

0180-26495-4 Vol. IV

D180-26495-5 Vol. V

D180-26495-6

0180-26495-7

0180-26495-8

Executive Summary

Requirements (NASA CR-160944)

SOC System Oefinition Report

SOC System Analysis Report (2 volumes)

Data Book {Limited Distribution}

(Reserved)

Space Operations Center Technology Identification

Support Study, Final Report

Final Briefing Book

For convenience to the reader, a complete listing of all of the known Space Opera-

tions Center documentation is included in the Reference section of each document.

This includes NASA, Boeing, and Rockwell documentation.

ii

D180-26~95-4

KEY TEAM MEMBERS

?ubject;

SOC Technical Manager

SOC T(~chnology Study Manager

System Design

Electrical Power

ECLSS/EVA

Communications & Tracking

Structures/Dynamic Control

Sta.b/Control

Propuls ion/Propellant Storage

Subsystem Interface

Programma tics

Software/Processing

Config. Design/Docking &

Berthing

Health Maintenance Facility

Crew Habitat

Operations

Space Construction Facility

Flight Support Facility

Crew Operations

Orbital Altitude

Opera tions Concepts/

Requirements

Transporta tion

JSC-Managernent Team

S. H. Nassiff

R. Kennedy

S. H. Nassiff

L. Murgia

D. Thompson

R. Dietz

R. Wren

J. Bigham

D. Kendrick

L. Monford

R. Kennedy

E. Dalke

J. Jones

D. Nachtwey

M. Dalton

B. M. Wolfer

L. Jenkins

H. Patterson

M. Dalton

F. Garcia

B. Wolfer

B. Wolfer

iii

Contractor Team

G. R. Woodcock

R. L. Olson

G. R. Woodcock

S. W. Silverman

K. H. Miller

H. Brose (Ham Std)

G. Rannenberg (Ham Std)

R. Cushman (Ham Std)

F. M. Kim

R. M. Gates

J. H. Mason

G. R. Woodcock

M. A. Stowe

G. R. Woodcock

G. R. Woodcock

L. E. Silva G. L. Hadley

J. J. Olson

K. H. Miller

K. H. Miller

K. H. MiUer

K. H. Miller

K. H. Miller

G. R. Woodcock

K. H. Miller

K. H. Miller

G. R. Woodcock

K. H. Miller

K. H. Miller

G. R. Woodcock

Subject

Technology

Cost

DI80-26495-4

KEY TEAM MEMBERS (Continued)

JSC-Management Team

R. Kennedy

W. H. Whittington

iv

Contractor Team

E. A. Gustan

R. L. Olson

G. R. Woodcock

T. Mancuso

D180-26495-4

TABLE OF CONTENTS

FORWARD

KEY TEAM MEMBERS

ACRONYMS AND ABBREVIATIONS

Section 1.0

1.1

1.2

1.3

1.4

Section 2.0

Section 3.0

Section 4.0

Section 5.0

Section 6.0

Secti()n 7.0

Section 8.0

Section 9.0

Section 10.0

Section 11.0

Section 12.0

Section 13.0

Section 14.0

Section 15.0

Section 16.0

Section 17.0

Section 18.0

BOOK 1 OF 2 --- -=-:~.:.....;:.....:::..;;.~--------,-,----"---INTRODUCTION AND BACKGROUND

Background

Program Objectives and Assumpti()ns

Study Purpose and Scope

Document Purpose

SUMMARY OF TRADES AND OPTIONS

MISSION ANALYSIS

CONSTRUCTION ANALYSIS

FLIGHT SUPPORT FACILITY ANALYSIS

SATELLITE SUPPORT FACILITY ANALYSIS

CREW CONSIDERATIONS

REFERENCES

BOOK 2 OF 2

ORBIT ALTITUDE SELECTION ANALYSIS

ELECTRICAL POWER SYSTEM ANALYSIS

ENVIRONMENT AL CONTROL/LIFE SUPPORT SYSTEM ANALYSIS

COMMUNICATIONS AND TRACKING SYSTEM ANALYSIS

STRUCTURAL DYNAMICS ANALYSIS

CONTROL DYNAMICS ANALYSIS

SOFTW ARE AND DATA PROCESSING ANALYSIS

PROPULSION AND PROPELLANTS ANALYSIS

SUBSYSTEM INTERRELATIONSHIPS ANALYSIS

SYSTEM DESIGN/OPERATIONS ANALYSIS

PROGRAMMA TICS AND DEVELOPMENT ANALYSIS

REFERENCES

v

AAP

AC

ADM

AM

APC

APSM

ACS

ARS

ASE

BIT

HITE

CAMS

C&D

C&W

CCA

CCC

eEl

CER

CF

CMG

CMD

CMDS

CO2 CPU

CRT

dB

DC

DCM

OOT&E

DOD, DoD

OT

DM

OMS

DSCS

D180-26495-4

LIST OF ACRONYMS AND ABBREVIA nONS

Air lock Adapter Plate

Alternating Current

Adaptive Delta Modulation

Air lock Module

Adaptive Predictive Coders

Automated Power Systems Management

Attitude Control System

Air Revitalization System

Airborn Support Equipment

Built in Test

Built in Test Equipment

Continuous Atmosphere Monitoring System

Controls and Displays

Caution and Warning

Communications Carrier Assembly

Contaminant Control Cartrige

Critical End Item

Cost Estimating Relationships

Construction Facility

Control Moment Gyro

Command

Commands

Carbon Dioxide

Computer Processor Units

Cathode Ray Tube

Decibles

Direct Current

Display and Control Module

Design, Development, Test, and Evaluation

Department of Defense

Docking Tunnel

Docking Module

Data Management System

Defense Satellite Communications System

vi

ECLSS

EDC

EEH

EIRP

EMI

EMU

EPS

EVA

EVC

EVVA

FM

FMEA

ftc

FSF

FSS

GN&C

GEO

GHZ

GPS

GSE

GSTDN

GFE

GTV

HLL

HLLV

HM

HMF

HPA

HUT

Hz ICD

IDB

IOC

IR

D 180-26495-4

LIST OF ACRONYMS AND ABBREVIA nONS (Cont.)

Environmental Control/Ufe Support System

Electrochemical Depolarized CO2 Concentrator

EMU Electrical Harness

Effective Isotropic Radiated Power

Electromagnetic Interference

Extravehicular Mobility Unit

Electrical Power System

Extravehicular Activity

EVA Communications System

E VA Visor Assembly

Flow Meter

Failure Mode and Effects Analysis

Foot candles

Flight Support Facility

Fluid Storage System

Guidance, Navigation and Control

Geosynchronous Earth Orbit

Gigahertz

Global Positioning System

Ground Support ~quipment

Ground Satellite Tracking and Data Network

Government Furnished Equipment

Ground Test Vehicle

High Level Language

Heavy Uft Launch Vehicle

Habitat Module

Health Maintenance Facility

Handling and Positioning Aide

Hard Upper Torso

Hertz (cycles per second)

Interface Control Document

Insert Dr ink Bag

Initial Operating Capability

Infrared

vii

IVA

J5C

KBPS

KM, Km

KSC

Ibm

LCD

LCVG

LED

LEO

LiOH

LM

LPC

LRU

LSS

LTA

LV

Ix

MBA

rnbps

MHz

MMU

MM-Wave

MOTV

MRWS

MSFN

N/A

NBS

NSA

N

NiCd

NiH2

Nm, nm

N/m2

D 180-26495-4

LIST OF ACRONYMS AND ABBREVIA nONS (Coot.)

Intravehicular Activity

Johnson Space Center

Kilo Bits Per Second

Kilometers

Kennedy Space Center

Pounds Mass

Liquid Crystal Display

Liquid Cooling and Ventilation Garment

Light Emitting Diode

Low Earth Orbit

Lithium Hydroxide

Logistics Module

Linear Predictive Coders

Lowest Replaceable Unit

Life Support System

Lower Torso Assembly

Launch Vehicle

Lumens

Multibeam Antenna

Megabits per second

Megahertz

Manned Maneuvering Unit

Millimeter wave

Manned Orbit Transfer Vehicle

Manned Remote Work Station

Manned Space Flight Network

Not Applicable

National Bureau of Standards

National Security Agency

Newton

Nickel Cadmium

Nickel Hydrogen

Nautical miles

Newtons per meter squared

viii

OBS

OCS

OMS

OTV

PCM

PCM

PEP

PIDA

P/L PLSS

PM

ppm

PRS

PSID RCS

REM

RF

RFI

RMS

RPM

SAF

SAWD

scfm

SCS

SCU

SEPS

SF

SM

SOC

SOP

SSA

SSP

SSTS

STAR

D 180-26495-4

LIST OF ACRONYMS AND ABBREVIATIONS (Coot.)

Operational Bioinstrumentation System

Onboard Checkout System

Orbital Maneuvering System

Orbital Transfer Vehicle

Pulse Code Modulation

Parametric Cost Model

Power Extension Package

Payload Installation and Deployment Apparatus

Payload

Portable Life Support System

Power Module

Parts per Million

Personnel Rescue System

Pounds per Square Inch Differential

Reaction Control System

Reoentgen Equivalent Man

Radio Frequency

Radio Frequency Interference

Remote Manipulator System

Revolutions Per Minute

Systems Assembly Facility

Solid Amine Water Desorbed

Standard Cubic Feet per Minute

Stability and Control System

Service and Cooling Umbilical

Solar Electric Propulsion System

Storage Facility

Service Module

Space Operations Center

Secondary Oxygen Pack

Space Suit Assembly

Space Station Prototype

Space Shuttle Transportation System

Shuttle Turnaround Analysis Report

ix

STDN

STE

TBD

TDRSS

TFU

TGA

TIMES

TLM

TM

TT

TV

UCD

VCD

VDC

WBS

WMS

DI80-26495-4

LIST OF ACRONYMS AND ABBREVIA nONS (eont.)

Spaceflight Tracking and Data Network

Standard Test Equipment

To Be Determined

Tracking and Data Relay Satellite System

Theoretical First Unit

Trace Gas Analyzer

Thermoelectric Integrated Membrane Evaporation System

Telemetry

Telemetry

Turn tab Ie /Tilttab Ie

Television

Urine Collection Device

Vapor Compression Distallation

Volts Direct Current

Work Breakdown Structure

Waste Management System

x

8.0 ORBIT All TITUDE SELECT1[ON ANALYSIS

8.1 ATMOSPHERE MODEL. 8-1

8.2 ORBIT DELAY TIME. . 8-1

8.3 PROPELLANT CONSUMPTION

8. 11 ORBIT ALTITUDE SELECTION 8-7

01S0-26495-4

S.o ORBIT ALTITUDE SELECTION ANALYSIS

B.1 ATMOSPHERE MODEL

Four atmosphere illodels were useLI in derivin\:j the orbit decay data, see Fiyure

Cl-I. The nominal model is the u.s. Standard Atmosphere, 1976. The other three II10deis were ~enerated vii!. tile quick-look density mOllel in At->t->e.ndix B of

NASA docUlilt:nt SP-S021, Models of Earth's Atmosphere (9U to 2~UU kill), for a

latitude of 00 This 1II0del calculates an exospheric temperature. i\ data

tab 1 e is then used to obta i n the lo!:) of the atmospheri c dens ity for the

desired altitude(s) for the calculated temperature. The NASA Neut ral dnll

Short Time Maximulll Models use val ues suggested for space shuttle studies. Hie

NASA Neutrdl Model is a high solar activity model, wittl a value of 23U for the

Inean lO,,] Cill solar flux and a geomagnetic index (Ap) of 20.3. The dSSUIlied

local lillie is U~UU hours. nil! Short Time IViaximUlIi IVlodel uses a h.l.7 crn solar flux of 23U, a geomagnetic index of 400, and a local time of 140U rlOurs.

These conditions would occur only for a tilile of 12 to 36 hours durill~ an

extreilleiy 1 ar~e Illagnet ic stonn. The IVli nillluHl l"Iodel uses fi gures of o~uu for

the IOCdl tillie, 7;).3 for the lU.7 CIII solar flux, and 10.9 for the ~eolllagnetic

index. The sol ar f1 ux and geomagnetic index fi gures are the 97.7 percent i Ie

fi~urt.:s for June 1%7 from tile IVldrshal"1 Space FI ight Center predictions.

The "NASA neutral" is considered to be the ~wrst long-term or continuous case

af)plicable to the 90-day resuf)ply cycle. The short-time maxilflufll \1ill be used

to estdbl ish thrust levels needed for control authority in all situations.

8.2 ORBIT DECAY TIME

Altitude Selection is based on maintainin\:j a minimum orbil: decdy tilfle of ~ll

days if no orbit maintenance occurs.

,--Tile velocity \JdS cdlculateu using V =; M x WUU Wflere M is the ~rdvitdtiollal rr

8-1

D 180-26495"-4

10'""12 12 ..

5xl0'"" - ....

21m 250 300 350 400 450 ALTITUDE IN KILOMETERS

Figure 8-1. Atmosphere Density Models

8-2

0180-26495-4

coefficiellL, eljuul LO 3~(5,6U1.2 krn3/secL, r is the rddius of the orbit, 1I1(~asured frolll the center of the Earth in kilometers, and the lOUO is a

cOllversion factor to yet the results in meters/sec.

COApV 2

Dray was calculated by f = 2 " where Co is the dray coefficient, A is the fronta 1 area in meters, pis the atmospheri c dens i ty in ky/m3 dno Vis the

velocity.

Mr COAp (8.64 x 107) The decilY rate \~as obtdined from the formula D = M ~~here CO' A, p dre as previously designated, M is the mass of the SOC in k~, dnd d.G4 x lU7 is the conversion factor to yet the results in kmjddY.

The decay tiliie was calculated from Q,o. =: Q~-l + (~ + ~) H Hhere l\ is the dl!Cay rilte at altitude x and H is the differencBA"Ot Bte t~JO altitudes. The SOC dldraneristies used in this set of calculations are Co = 3.0, A = 300m2, M = lUU,OUOky, and I = 23U sec. sp

Fi~ure 8-2 ShOHS the dltitude requirelnent as a function of atmosphere IIlodel.

The fi~ure is based on a lIledn CdA of 1800 square meters and a SOC mass of IUU

tonnes. Since the dctUd 1 SUC lIIass wi 11 pr'obab ly exceed th is fi yure, the curve

is slightly conservative.

F i yure 8-3 sho~~s atlflospheri c dray versus a 1t i tude for the four IIlode 1 sand

Figure 8-4 shows the orbit decay ranye. (Fiyure 8-2 was derived by numericcll

i nte'Jrat i on of Fi yure 8-4.)

8.3 PROPELLANT CONSUMPTION

Fi~ure 6-b shows the prOPellant consumption \~ith the decay altitude lilll"iL

superililposed. Propell ant consumption was based on the use of Illonopropell ant

IlytlruLine ilt a specific impulse of L3U sel.:.

The jJrope 11 ant usaye was equal to 86400f 9 Isp

8-3

where f is the dray force in

z : Hllil =: ;; 50 :::~. I--

>--< tj 10 o 5-I--

ill 0: o

D180-26495-4

., !

NOMINAL

I

1..1-L1.l 2013 2513 3013 3513 41313 450

ALTITUDE IN KILOMETERS

Figure 8-2. Orbit Decay Time lIS. Altitude

8-4

D 180-26495-4

1I:l0 ~-. 1 so : ,,~... ,

~ ,,~'-...'

f !

"I

" ,........... l

',"":::::--.... ; I , ..... '-.... .... J........; . I ' --..... : .',' "~'''''''''''''_i ' . ..... , ................ ; 1 .... , ; ............ : -...;-. I

, "'-" ~!-.. ... -.. '1- ... _--~ ".....,......:-_.;! - .. - . ......... --.. .... ...... ... ~: gj .S:: "i -..' --_ '~SHORT " --., ----. --.. ...... ~ U : ".... ... I : TIME .... ~-. ! .... .... i -.. .... ~ . MAXIMUM f.5 I ..... ' --...~ :I: .1 :~.....: --'''SNASA 5; ~: . ~ 1 .......... "' : .. : NEUTRAL ! .. os r. ' ! ... , ~ t~.. i! . ! ... .... ... ... ... ,. ~ NOMlHAL

, I : l' I .... ' .. 01 ~:: 1 , ' l' '. . '1' " .... '1 ... '" . .~ -. I, I .. ... ' .

~05 ELl.LLLLLL ,J,.LL,!. "J.LL:J .. LJ .. _LL .... LL -1. LL.L1J.::1;iMINlMUM 200 250 300 3S0400 450

ALTITUDE IN KILOMETERS

Figure 8-3. Atmospheric Drag vs. Altitude

"--r'" .. HlO ~ '. 50 .. ,,~.... ,

::.,~.

z ....

" ..... ~" ' ........ ,,:.-" .....,; .......................... ........ '...... ..... ............ '""to ..... ---..

...... ......-.. ..... .... ........ -...

' ........ ~ .... -....:. ... _--_. "-..... ...... . . ~ -''' ..... - -...... ':.: SH(Rf ........ . ............ ... ... __ = TIME

.... ..... -.. - _ . : MAXlMUN .......... : ' - -.......... -

....... -_.NASA ........ NEUTRAL

5 =--

t; . 1 m .. 05 -o

.... .... ........ ........ NOMlHAL

.... .... ........

...... '" .: 1 l..l I LI'j MINlMUN

.. 01 .--, i I .,105 LL!J J LJ ,1...1 l.L.LJ LLL.L.LL .. .1 l.lLL[

200 250 300 350 400 4S0 ALTITUDE IN KILOMETERS

Figure 8-4. Orbit Decay Rate vs. Altitude

8-5

D180-26495-4

IIOC39I

r"-' ".--" . -." .. -.-~

t; 10 . , =< 5 ...... - .. ...J r-' .J W n. 0 a:: n.

"- ... .... .... .... ' ...... NOMINAL ...

.5 ~ r- -~-

1 LLI .1 i

. i: 1 .. ..;. . "-.; ~.:.. ...... --, . .L .. I .. LL I .. L_L .. Li.Li.l_L--L. 1LL1._L1..1 . .1-.L .. LL.L:'l=l~ MINIMUM

200 250 300 350 400 450 ALTITUDE IN KILOMETERS

Figure 8-5. Propellant Usage vs. Altitude

8-6

D1()U-2b4~!:>-4

newtons, y is the dcceleration of grdvity, 9.81 m/secL , lsp is tile specific

impulse of the IIlotors, and 8G,400 is the conversion factor to yet the resuits in k.y/ddY.

A resu(Jply reL\uirement of LL k'::J/day has been defined based on tilis curve, with

2k'::J/day added for atmosphere makeup (from hydrazi ne decompos it i on) and a lU,o

lIlar~jin on the NASA neutl~al atmosphere point. The nominal resul-lply reL\uirelfient willi be somewhat less.

8.4 ORBIT ALTITUDE SELECTION

The selected orbit altitude is 37U km, as this is the altitude the shuttle Cdn

reach ~Jithout OMS kits. This provides the maximum payload bay lenyth

capabil ity. The mission II10dei analyses show tllisto be extremely ililP0rL.dllL.

During per"iods of hi'::Jh solar activity, the altitude will be raised to 4UO kill.

There are several operational options available to deal with the possibility

of needing full shuttle payload bay when the SOC is above J70 km.

The Y'esuPf.lly requireillent tldS been set at L2UU kg for' lUU days in sizin~ the

10!;jistics module. This requires only one ring of six 1.12 HI ,4411) tanks on tile "logistics Inodule.

&-7

9.0 ELECTRICAL POWER SYSTEM ANALYSIS

9.1 INTRODUCTION. . . . . . . . . . . . . . . . . . . . . 9-1

9.2 ELECTRICAL LOADS DATA 9-1

D18U-l6495-4

9.0 ELECTRICAL POWER SYSTEM ANALYSIS

9.1 INTRODUCTION

Thl~ bu I k of the e I ectri Cd I power systeill ana lys is datd is inc 1 uded in the SUL

System Uefinition Report (Boeiny - 19) under WBS 1.2.2.1.7 so it is not

repeated herein. Section 9.2 gives the lower-level electrit:al 10dd tables.

Weiyht pendlty calculations are found in the Data Book (Boeing - ill.

9.2 ELECTRICAL LOADS DATA

Table 9-1 ::liVeS tile eleccrical "load sUllllllary. Tile life support e4uipment IOdds

~vere taken frolll Table B in WBS 1.2.1.1.13 in the SOC System Definition f{eport

(Boein,::)-19). The other subsystem electrical loads are detailed in Table l!::!-i.

9-1

SOC-1117

Table 9-1. Electrical Load Summary

REFERENCE CONfiGURATION INTERMITTENT LOAD

SUNLIGHT OCCULTED POWER

AC DC AC DC AC DC

LifE SUPPORT 19W 10,209W 6,565W 2,61 OW 3,85OW 3,75OW COMMUNICATIONSffELEMETRY 9,370W 9,270W DATA MANAGEMENT SYSTEM 1,OOOW 1,OOOW PROPULSION SYSTEM 200W 200\'\I t::J THERMAL CONTROL SYSTEM 2,00OW -2,00OW 00

CONTROL SYSTEM* 250W 250W 0 I N .. ICAL POWER 0\ ~

LOADS 12,50OW 4,500W \0 VI \J:) BATTERY RECHARGE 29,900W

I i ~ N

TOTALS 19,91gw 51,429W 19.365W 19,83OW 3,85OW 3,75OW

* 1 KW STARTUP - 6 HR {CAN BE SUPPORTED BY LOAD DIVERSITY)

[ .! ",,,,,,;,~ l;;;;~'" I I I I > 0 (.) 112 SOC CONFIG CONFIG 0 Z w w w 0 . .-,~-- 0 C) l- s: I- ~ I- w Z a: a: CC-n32 ::t: i= w C) ;- Z C) ..J S1 ..J 12 ...J 0 ::IE w :::; ::l ...J ::l ...J ::l DIST. ...J ::IE w 0 Z (.) Z

~ Z g FROM

10.0 EeLS AND EVA/IVA STUDIES

10.1 VENTILATION CONCEPTS STUDIES. . . . . . . . . . . . .. 10-1

10.1.1 Forced Convection Level Required to Simulate Free Convection 10-1

10.1.2 Use of Ventilation Direction to "Simulate" Gravity. 10-1

10.1.3 Flow Required for Cabin Heat Rejection . . . . . 10-2

10.1.4 Clean Air Supply Ducts vs. "Dirty" Air Return Ducts 10-3

10.1.5 Ventilation During SOC Buildup . . . . . . . . 10--5

10.1.6 Summary Description of the Selected Ventilation Concept. 10-6

10.2 EMERGENCY PERFORMANCE LEVEL DEFINITIONS. 10-7

10.3 CABIN PRESSURE ASSESSMENT . . . . . . . . . . . . . 10-10

. 10-11 10.3.1 Factors Influencing Selection of Cabin Pressure ....

10.3.2 The Effects of Selected Cabin Pressure on ECLS System

Com ponents II............... 10-16 10.3.3 Conclusions Regarding Selection of Cabin Pressure. . 10-18

10.4 REDUNDANCY PHILOSOPHY .... 10-22

10.5 MAINTENANCE CONCEPT DEFINITION 10-24

10.6 EVALUATION OF SELECTABLE CABIN TEMPERATURE AND CABIN

TEMPERA TURE BAND EXPANSION . . . .

10.7 ECLS SUBSYSTEM SELECTION RATIONALE ...

10.7.1

10.7.2

CO2 Removal Subsystem Concept Selection.

Wastewater Processing Concept Selection ..

10-33

. . . . 10-36

. . 10-36

10-50

D180-26495-4

10.0 EeLS AND EVA/IVA STUDIES

10.1 VENTILATION CONCEPTS STUDIES

10.1.1 Forced Convection Level Reguir!d To Simulate Free Convection

Manis physiology is equipped to reject body heat and moisture without wind across his body, providing he is in the earth's one gravity atmosphere. The effect of gravity is to induce a quantity of convective heat transfer and air mass transfer, driven by the change of density occurring near the surface of the body. This convective force is not present at zero gravity, making necessary an artificially induced convective ventilation in order to simu-late the free convection which is lost. This phenomenon has been evaluated and lived with in all previoUls spacecraft, and a fan induced average velocity of 25 feet/min has evolved as the accept-ed ventilation design value for spacecraft.

10.1.2 Use Of Ventilation Direction To "Simulate" Gravitl

Man's physiology and geometry are configured to reduce the likeli-hood of eye damage or choking from loose objects, such as some-th'ing dropped while eating or dropped from the hands. The eyes and mouth are "up", and things nOY'mally fall the other way, "down". In a zero-gravity environment this characteristic of getting things to fall dO~'1n may be partially simulated by util-iz~ing a ventilation system which has 'its cabin airflow descend from ceiling to floor. This concept has been selected for SOC.

However, it is not practical that the 25 feet/min ventilation vel~ocity of the previous paragraph be entirely made up by this dov~nward flow. Every air jet or anemostat is in effect the pri;mary nozzle of an ejector which induces many ,m,ultiples of secondary flow into its flow pattern. This in turn results in a circulation pattern where flow is concentrated in a downward, direction under the anemostat, proceeds down toward.the floor, and then circles back toward the anemostat and around again to rejoin the downward flow. The net flow is downward. but locally there is increased velocity in the down direction under the outlets, and in

10-1

0180-26495-4

an up direction between the outlets. For SOC, the primary flow of

the vent supply anemostats ;s sized by the flow required to pass

through the heat rejection heat exchangers in order to maintain

cabin temperature, as discussed in the next paragraph. If a

downward velocity of 25 ft/min were incorporated, the power con-

~umption of the ventilation system would be about 3 to 4 times the

selected baseline power requirement.

10.1.3 Flow Required for Cabin Heat Rejection

There are two choices in selecting the quantity of airflow which

is to be cooled in the heat rejection heat exchangers. One way is

to use 40F coolant fluid through the heat exchanger and calculate

the airflow required to transfer the heat load. The 40F value is

selected as the lowest feasible temperature to avoid freezing in

the coolant water to freon heat exchanger. In thi s case, there

would be condensation in the cooling heat exchangers, due to the

coolant being below the desired cabin air dew point. The moisture

thus removed would be collected at each heat exchanger and pumped

to the water processing system. The relative humidity of the air

leaving these heat exchangers would be excessive, since the air

would be essentially saturated. Also, when this method of select-

ing the airflow to be cooled in the heat rejection heat exchangers

is utilized, the resulting airflow is too low for use as the

primary anemostat flow to provide the required 25 ft/min local

velocity in the cabin. Additional cabin air circulating fans

would be necessary to raise the cabin air velocity to the required

level. The above method of selecting airflow for the heat rejec-

tion heat exchangers was not selected because of the complexity of

removing moisture at each heat exchanger, and the complexity of

additional circulating cabin air fans.

The other way in which cabin airflow through the heat rejection

heat exchangers can be sel ected was used in the SOC basel i ne

system of this report. In this case, the coolant fluid is con-

trolled to 55F entering the heat exchanger, rather than 40F as

described in the previous paragraph. This prevents condensation

of moisture present in the cabin air by keeping metal temperatures

10-2

D180-26495-4

over the dew point, and eliminates the problems of separating,

collecting, and pumping water, at each heat exchanger, as wel"' as

eliminating the possibility of fog generation at the cabin supply

anemostats. It also provides adequate primary flow in the cabin

supply anemostats to efficiently provide the desired cabin air

velocity.

Both of the above methods of selecting cabin air heat rejection

airflow were evaluated as part of the Space Station Prototype

(SSP) program, and the second method was selected. There is no

significant difference in requirements for SOC which would indi-

cate that the selected concept for SSP should not be the preferred

concept for SOC.

A schemat"ic of this ventilation concept, as selected for SOC, is

shown on Figure 10-1. Note on the figure that the principle of

IIdo~lnward" net airflow of Section 10.1.2 is accomplished by air

supply anemostats in the ceiling, supplemented by local adjustable

anemostats in accordance with the detailed floor plan (such floor

plan details will become evident later in the SOC development).

These supply anemostats are fed from a common plenum over the

ceiling, which in turn is fed by the sum of flow leaving the

temperature control heat exchangers plus the flow leaving the air

revitalization packs. The tE!rm "air revitalization pack" is used

here to describe the equipment group which includes the functions

of removal of humidity, CO 2 ' odor, and trace contaminants. De-

tails of this equipment group are summarized in WBS 1.2.1.1.13.2

in Boeing-19. The term "ventilation and temperature control" pack

is used to describe the equipment group which includes ventilating

airflow fans, particulate filtration, heat rejection heat exchan-

gers, and appropriate sound suppression baffling, as described in

WBS 1.2.1.1.13.1 in Boeing-19.

10.1.4 "Clean Air Supply Ducts vs. "Dirty" Air Return Ducts"

Another choice in design of the ventilation system is the way in

which odor and moisture sources are handled. A supply of "fresh"

air could be specially ducted to the toilet area, for example, or

10-3

0180-26495-4

a r'eturn duct carrying IIdirty" air from the odor source could be utilized. The only reason for considering the "fresh" duct option is that d shorter, reduced volume duct could possibly be selected for baseline, in spite of its potentially larger duct volume. Most

of the increased odor and moisture control duct volume will be in the overhead plenum on SOC, and it is presumed that SOC is a second generation spacecraft where such personal anemities as odor control should be! considered. No trade-off on quality of living is pos-sible, but the selected concept of controlling local odor and humidity sources appear to be common sense. In any case, the volume difference between these duct options is not a major matter, and furthermore an exact calculation of the difference in vol ume

between the two concepts is impossible.

10.1.5 Ventilation During SOC Buildup

It is not currently envisioned that thE! SOC will be permanently

inhabited until all baseline modules are in place. However, during the buildup sequence the crew may pressurize and enter the service module from the Shuttle. The service module will have its own power. Limited heat rejection will be provided by the battery and power conditioning equipment radiator. Humidity, and CO 2 control can be provided by using a snorkel line from the Shuttle air revita-

lization system. This capability (48 cfm) is a standard capability of the Shuttle for use with the Space lab. Some thermal control is also provided by this air f'low. Only one of the service module ventilation fans would be needed to provide minimum acceptable air mixing and air velocity for cooling.

After the first habitat module is in place the half SOC config-

uration will have an operational EClS system. No air flow mixing between the Shuttle and the habitat is required. The Shuttle must remain attached to provide adequate safety in case the habitat must be evacuated.

10-5

0180-26495-4

10.1.6 Summary Description Of The Selected Ventilation Concepts

The selected ventilation concept was shown schematically on Figure

10-1. Local anemostat detail s will not become apparent unti 1 an

actual design phase.

One major air recirculation path shown on Figure 10-1 consists of

return grills in the floor leading to an under floor plenum, with

two large vertical return ducts (approximately one square foot

each) connecting the under floor plenum with the ventilation and

temperature control packs in the overhead plenum. Each of these

ducts supplies one IIdouble ll vent pack in the overhead plenum.

These IIdouble ll packs each contain two independent temperature

control systems as are described in detail in WBS 1.2.1.1.13.1 in

Boeing-19. A half of one of these double ventilation and temper-

ature control packs is also utilized independently at each end of

the service module, as shown on Figure 10-1. In other words there

are four ventilation half packs in each habitat module (in two double units), and two other half packs in each service module,

making a total of six per half SOC or twelve per full SOC. This

a p pro a c h pro v ide sma x i mum co mm 0 n ali t y 0 f h a r d war e and imp r 0 v e s

reliability compared to the option where separate vent packs would

be sized for the habitat module and the service module.

Another air circulation path shown on Figure 10-1 consists of the

flow of dirty. odorous, or wet air taken from areas of contami-

nation, and ducting to two air revitalization packs arranged in

series in the overhead plenum. The suggested source areas shown on

Figure 10-1 include the shower, toilet, suit storage area, and the

remote end of the service module for contaminant control in the

service module. Roughly 5 percent of the total habitat module

supply airflow passes through the contaminant removal packs, so

that it takes 76 minutes for them to pass an airflow equal to that

of the entire cabin volume.

The total habitat module cabin supply airflow of 2440 CFM enters

the cabin through anemostats placed several feet apart at inter-

vals in the ceiling, and through adjustable anemostats placed as

10-6

D180-26495-4

dictated by the final design. The velocity of these cabin supply

airflow nozzles will generate an induced secondary airflow approx-

imately 4.5 times the primary flow. The resultant total flow is

adequate to provide an induced flow of 25 ft/min local velocity.

of '''hich the average up flow is 25 ft/rnin and the average down

flow is 32 ft/min, resulting in a net down flow down of 7 ft/rnin.

The induced flow performance of the primary airflow system select-

ed for SOC is based on studies and development activities per-

formed as part of the SSP program.

Figure 10-1 considers the baseline floor plan for the habitat

mod u 1 e con sis tin 9 0 f a sin g 1 e 1 0 n g i t u din a 1 floor. Ot her 0 p t i () n a 1

floor plans include "baloney slice" floors for at least a portion

of the module. As far as the ventilation concept of this report

is concerned, it is recommended that the basic concept resented be

used regardless of the floor plan selected. Obviously the imple-

mentation of the ventilation system is easiest with a single

longitudinal floor, and this is certainly one of the advantages of

such a floor plan. If other factors should predominate, and a

baloney slice floor plan is adopted, the ventilation duct system

becomes more complex and difficult to visualize. The basic flow

pattern from ceiling to floor should be preserved where feasible.

10.2 EMERGENCY PERFORMANCE LEVEL DEFINITIONS

A SOC requirement imposed by NASA in document NASA-6 directed a

fail operational/fail safe design criteria. This criteria, with

minor excE~ption has been retained. This minor exception is that

the "operational level" has been assumed a level which provide an

acceptable performance for a 90 day period. In order to meet this

safety requirement it is believed that no single failure of ECLS

equipment shall force abandonment of a habitat module. This

establ ishes the basic requirement for dual radiators, dual radi-

ator transport loops, and dual atmosphere supply and processing.

Each of these dual systems is not capable by itself to maintain

10-7

0180-26495-4

the excellent environment referred to as the lIoperationalll per-

formance level, since the result would be an unnecessarily large

vehicle penalty. Instead, it will take both of the dual systems

operating together to produce the lIoperationalll level. One of the

dual systems operating will produce a 1190 day acceptable ll envi-

ronment. By these steps of logic it is possible to design a

system which is fail operational (acceptable)/fail safe and which

imposes a minimum penalty to the vehicle. The performance level

capability of the EelS system is summarized on Table 10-1.

Referring to Table 10-1, note that the lIoperationalll performance,

which results without system failure and with the normal habitat

modul e crew compl ement of 4, is what coul d be descri bed as an

excellent environment. This performance can be maintained in

steady state for all mission activities of the 4 crew members.

More than 4 crew members in that habitat module will not cause a

noticeable reduction in the quality of the environment for rela-

tively short duration activities such as meals or meetings.

Neither will there be a noticeable reduction in environment qual-

ity for longer periods if the activity level is low, such as 8 men

sleeping. However, more than 4 crew members in the habitat module

continually, with full activity level, wash, shower, cooking, etc,

can cause the performance level to approach the 1190 day, acceptable ll

level. The table shows that this same 1190 day acceptable ll perfor-

mance level can be maintained with a 4 man complement after a worst

single non-maintainable failure in that module. This capability

enables the system to meet the fail operational criteria of not

having to abandon the habitat module with a worst single failure.

If the failure such as a major fire or major breach of the cabin

wall, has forced abandonment of a habitat module the entire crew

will be in the remaining habitat. In this situation the reduced

level of performance capability is still 1190 day acceptable ll As

shown on Table 10-1, 8 man capability is provided for at least 300

10-8

..... a I \0

TABLE 10-1

ECLS PERFORMANCE LEVEL REQUIREMENTS

Parameter

CO2 Partial Pressure

Temperature

** Dew i'~}int Temperature

ventilation

Wash Water

*** 02 Partial Pressure

Total Pressure

Trace Oontaminants

Units

mmHg

OF

OF

ft/min

1bj."roan day

psia

psia

Maximun number of crew per habitat without failure in each habitat

Maximum number of crew per habitat with worst single non-maintainable failure in that habitat

:190 "(perational" Acceptable"

3.8 max 7.6 max.

65-75 60-85

40-60 35-70

15-40 10-100

40 min 20 min

2.6 or 3.1 2.4-3.8

10.0 or 14.7 10.0-14.7

**** **** 24 hr. indo stld. 8 hr. indo stld.

4 8

NA* 4

;;300 Hour Emergency"

12 max

60-90

30-75

5-200

o

2.3-3.9

10.0-14.7

**** 8 hr. irrl. st'd. 12

8

*Acceptable level is adaquate to meet a "fail operational" reliability criteria.

**In no case shall relative humidities exceed the range of 25-75%.

***In no case shall the 02 partial pressure exceed 26.9% or be below 2.3 psia.

****hr. indo stld. = hour industrial standard

o -co o I

N 0'1 ~ ~

0180-26495-4

hours at the "300 hour emergency" level with a single worst non-maintainable failure after the crew has abandoned one habitat. It should be noted that a single habitat has a 12 man capability at the levels shown in the 300 hour emergency column without a time limit if no EelS equipment has failed.

The design point used to size any particular EClS subsystem ;s generally found in one of the columns and that subsystem will exceed the required performance of the other two columns. There al so are subsystems which exceed the fail operational (accept-able)/fail safe design criteria because of redundance dictated by the requirement that no single non-maintainable failure shall cause evacuation of a habitat.

The performance levels listed on Table 10-1 were arrived at by establishing the lowest performance level that could be tolerated for the respective continuous 90 day and 300 hour day time per-i 0 d s

The specific values are a result of years of study. Information from actual flight experience (Apollo, Skylab, Gemini, etc.) and space station study programs 1 ike the Space Station Prototype (SSP) have been used to define the values shown. These values also have been discussed and reviewed with NASA/JSC during the conduct of this SOC study.

10.3 CABIN PRESSURE ASSESSMENT

The baseline SOC cabin pressure for purposes of this study is 14.7 pSia. but there are many factors which would favor the selection of a lower cabin pressure, as discussed in the following Section 10.3.1. However, selection of a lower cabin pressure has adverse e f f e c ton the s i z e , wei g h t , and power con sum p t i on. 0 f c e r t a i n portions of the EelS, as discussed in Section 10.3.2. An overall conclusion regarding the factors affecting cabin pressure selec-tion is presented in Section 10.3.3.

10-10

0180-26495-4

10.3.1 Factors Influencing Selection Of Cabin Pressure

~)xia - One factor to be considered in selecting cabin pressure

is avoidance of hypoxia, or reduced brain function, caused by low

partial pressure of oxygen. Due to the fact that the majority of

the earth1s inhabitants live below StOOD feet, the maximum con-tinuous altitude at which body functions are measurably affected

is not clear. A partial pressure of oxygen in the cabin corres-

ponding to 4.000 feet or less would certainly be desirable, and is

considered a requirement by NASA ("Medical Science Position on

Space Cabin and Suit Atmospheres" Position paper by NASA JSC/SD,

May 1980). Although an altitude as high as 8,000 feet equivalent OXYgen level is considered to be an acceptable level for commer-

cial aircraft pressurization.

Flammabil i"!'y - Another major factor to be considered in selecting

cabin pressure is flammability of materials. The fire danger is

relalted to the percent oxygen PTesent in the cabin atmosphere.

The normal sea level oxygen concentration is 21 percent, and certainly a concentration this low in the SOC cabin would be

desirable from a flammability standpoint. Only one major material

used in Shuttle, a silicon fiberglass line insulation, has failed

tom e e t fl a mm a b i 1 i t Y t est sat 35 per c e n t O2 , and t his mat e ria 1 will be replaced in later Shuttle vehicles. The cabin pressure

control tolerances of the current Shuttle result in a maximum

normal oxygen concentration of 23.8 percent 2 " A caution and warning light is set on Shuttle to trip at the 25.9 percent level with a 26.9 percent O2 absolute maximum level has been selected for SOC. These same levels are probably going to be inherited by

SOC as the flammability requirement. The relation between flam-

mability and cabin pressure is shown on Figure 10-2.

Eliminating Pre-breathe The pre-breathe period required to

prevent "bends" with the presently ava-ilable 4.1 psi suit is

approximately 4 hours pre-breathe with a 14.7 psia cabin, and approximately 2 hours pre-breathe with an 11 psia cabin. This lost time in pre-breathe can be eliminated by increasing suit

10-11

OXYGEN PARTIAL PRESSURE (PSIA) N N N N w w W

N ~

mr-~--1r-r~------~-------r~------~~------~~------4~ . . . . . .

"'OU'J ::00 trln n U'J ):' U'Jn \:XI c:):' 1-4 ::0 to ":rj z trll-4 1-4

0

Z G) "t! ......

n c: :;0 ..... co

O~ ::0 trJ 0

ZZ trl en ..... I

..... U'JtJ en N

0 H ..... c: trl 0'\

I tJ(f) :;0 10

.s=-..... 0 N trl"tl I trl

C eVA 0

\0

S;:~ N 1-4

()'1

~ PRACTICA 2 TOXICITY ,

8trl "t! ..... .s=-

en N t"' H 1-4 trl L" S OU'J ):' z .pIT CONST zc: 8 U'JH _~ RUCTION

8

..... w

trJ ~ .... 10 o -C g, 1-4

..... ~ ":rj t"' '8 ",. trJ

Z 8

-SEA LEVEL CABIN SHUTTLE

..... I U'I

DI80-26495-4

pressure or decreasing cabin pressure. For SOC where several

EVA's a day will be routine, eliminating pre-breathe is extremely

desirable. The relationship between suit pressure to avoid pre-

breathe and cabin pressure is also shown on Figure 10-2.

~gen Toxicity - The partial pressure of oxygen in a breathable

atmosphere must be limited to avoid toxic effects. The crew is

exposed continuously to the oxygen level in the module, as opposed

to only eight hours per day exposure in the EVA suit. Because of

this, the continuous oxygen cabin limit which can be tolerated in

the cabin is lower than the short duration EVA oxygen limit.

The upper limit of oxygen partial pressure select~d for Apollo and

Skylab cabins was 5 psia, and in the case of Skylab this was for

continuous use. There was some medical evidence of undesirable

oxygen toxicity in these programs, as reported in the literature

("Extravehicular Cre\'Iman Work System Study Program", Final Report,

Vol I I, Con s t r u c t ion , J u "' y 1 98.0 , Con t r act N A S 9 - 1 5 290 R. C

Wilde, Hamilton Standard). There has also been evidence of tox-

icity revealed in tests run since then, but there does not seem to

be a real consensus on the degree of seriousness of these observed

eff,ects. An oxygen concentration as high as 4 psia O2

partial

pressure could probably be tolerated continuously in the SOC

cabin, but this is a moot point because SOC will utilize a two gas

atmosphen~ making this high a PP0 2 unnecessary, as shown on the

left-hand vertical scale of Figure 10-2.

Oxygen toxicity during EVA is a different matter. First because

EVA will occur for an individual crew member for a maximum of

about 25 percent of his total in orbit time, and second because

the atmosphere in the suit will in all likelihood be pure oxygen.

The rei s so m e e v ide n c e t hat 8 . psi a pur e 0 xy g en pre s sur e i nth e suit will result in unacceptable toxicity effects, as described

in the literature (NADC-74241-40, "Physiological Responses to

Intermittl:nt Oxygen and Exercise Exposures", E. Hendler, NADC,

Warminster, PA, 1974). For eight hours a day, a 4 psia level is

generally accepted. The maximum allowable suit level of pure

10-13

0180-26495-4

oxygen level for EVA therefore, probably lies between 4 and 8 psia

but it is not a black or white matter, and considerable difference in tolerance between individuals undoubtedly occurs. A limit of 6 psia is logical since 4 psia is acceptable and 8 psia is not, but it is a tentative limit, not clearly defined. This tentative 6 psia suit pressure limit for pure oxygen is identified on Figure 10 - 2.

Weight of Stored Cabin Pressurization Gas The leakage flow

through any hole or leak in the vehicle pressure wall is directly proportional to cabin pressure. SOC cabin leakage is expected to be about 5 pounds of air a day. Another 5.3 pounds of air per day is expected to be lost in use of airlocks on an EVA day assuming

pump down to 2 psia for a 14.7 psia cabin. This total air loss is made up by oxygen produced from wastewater by electrolysis, and by nitrogen obtained from the decomposition of 9.3 lb/EVA day of hydrazine. Capability for one complete repressurization utilizing stored high pressure gas wei'ghs approximately 750 lb, plus tank-age. The weight of the above varies as follows with cabin pres-

sure:

Design Cabin Pressure

14.7 psia 11 ps i a 9 psia

Resupply Hydrazi ne Required For Nitrogen Makeup Per 90 Days

775 580 474

Stored Repressurization Gas, Including Tankage

1321 989 809

Resupply Water, Including Tankage Required For Oxygen Makeup Per 90 Days

270 202 166

Vehicle Mechanical Strength - Thickness of the SOC vehicle skin is dictated by the need for protection from meteorites and space junk. Reducing the vehicle cabin pressure would therefore not reduce skin weight.

10-14

0180-26495-4

Suit Considerations The EVA suit is presently qualified for a

nominal operating pressure in space of 4.1 psia. There is reason

to believe, however. that this could be raised to 4.5 psia without

significant difficulty. Beyond this, significant suit development

would be required. Certainly a 6 psi suit would be less complex

and more flexible than an 8 psi suit, and certainly it would cost

less to develop in terms of time and money. Although this latter

factor is not considered par'ticularly s"jgnificant in the overall

evaluation, the incredsed safety and dexterity resulting from a 6

psi suit, rather than an 8 psi suit, could be particularly impor-

tant on SOC where construction tasks and other functions of the

vehicle place such emphasis on EVA capability.

A major consideration of an 8 psi suit used at 8 psi gage at sea

level for training and development testing would require the

standby use of a hyperbaric chamber for safety in the event of a

suit pressurization failure. Rapid decompression may rupture

lungs putting air into a pleural cavity. The lung may collapse

and allow air into the blood. Decompression is essential to

reduce bubble size to reduce danger of air embolisum. The highest

suit pressure which does not require such a chamber for sea level

safety is approximately 6 psi.

A 6 psi suit is identified on Figure 10-2 as a IImost practical"

upp,er limit for suit construction purposes, but this is a judge-

ment call and not amendable to exact evaluation.

Adaptabil'ity of "Shelf Hardware!! to Shuttle and SOC - There is an

intangible benefit in utilizing a sea level cabin pressure in that

its h 0 u 1 d red u c e cos t by m a kin g i tea s i e r to uti 1 i z e co mm e r cia 1

items already developed for earth use. This intangible benefit no

doubt was a major factor in influencing Shuttle to be designed for

a 14.7 psia cabin. Unfortunately, the real value of a 14.7 psia

cabin in adapting shelf hardware is of less consequence than was

hoped. Only air cooled electrical equipment items are effected by

cabin pressure, and these are as much effected by the zero-gravity

effect of space as they are by cabin pressure level. The lack of

10-15

0180-26495-4

free convection cooling in space means that new fans will have to

be added anyway to most items which were free convection cooled on

earth. Once these fans are added, it is probably not a signifi-

cant cost increment to select them for the appropriate cabin

pressure. Figure 10-2 shows an 8,000 foot cabin pressure alti-

tude, which is typical of airline practice, for reference, but it

should be pointed out again that this altitude at zero-gravity

poses entirely different equipment cooling problems.

Commonality With Shuttle Cabin Pressure - The rationale being used

to select the final value of Shuttle cabin pressure becomes an

important consideration in sel ecting SOC cabin pressure as well.

since commonality between the two would be extremely desirable, if

not essential. Shuttle cabin pressure selection is not yet firm,

and in the event the final selected value for Shuttle differs from

that considered in this report, this SOC cabin pressure assessment

will require review and possible revision.

10.3.2 The Effect of Selected Cabin Pressure on EClS System

Components

The value of cabin pressure selected for design has many ramifi-

cations, as discussed in the previous section. One of these is

the fact that a lower cabin pressure makes rejection of heat from

the cabin air to the radiator coolant fluid more costly in terms

of system size. complexity, and power consumption. This is be-

cause cabin air is the first stage coolant for rejecting most of

the heat load generated in the cabin. This heat transfer is a

function of air mass flow, not CFM, and therefore reduced air

density increases the power needed to circulate the airflow re-

quired for heat transfer. This study considers a sea level cabin

pressure as baseline. If this hardware were built and developed,

and the cabin pressure were then reduced, the baseline heat rejec-

tion capability of the baseline EClS would degrade as shown on

Figure 10-3. This higher temperature may be undesirable so changes

to the system may have to be made to accommodate lower cabin pres-

sures. These changes need be made only in the components involved

in the cabin air temperature control and ventilation functions,

since :he other EelS systems components are unaffected.

10-16

...:l oet: z 0 H E-t oet: p:: t1l 0.. 0

:E ::> :E H x oet: :E

c... 0

t1l p:: ::> E-< oet: p:: t1l 0.. :E t1l E-t

z H OJ oet: U

85

80 ~

75

70 10

0180-26495-4

.....

~ BASELINE DESIGN POINT, ~

11

. r----~

- 12 13 CABIN PRESSURE (PSIA)

FIGURE 10-3

CABIN PRESSURE AFFECT ON CABIN TEMPERATURE

10-17

...

14 15

0180-26495-4

The simplest change which can be made to the EClS system to com-pensate for reduced cabin pressure would be to increase fan air handling capacity in order to maintain the design value of mass airflow, and accept the power and noise suppression penalty which would occur as a result. Figure 10-4 shows how fan power would increase to hold airflows, and therefore heat transfer, constant. Unfortunately, this solution of increasing fan size to accommodate

a lower cabin design pressure would add significantly to the electric load demand of the EelS. Figure 10-4 shows for example that an increase in power is required per full SOC from 3.6 kw to 6.5 kw, or a delta increase of 2.9 kw. in dropping cabin pressure from sea level to 11 psia. Battery weight needed to provide this

2.9 kw of power on the dark side would weigh 910 pounds. This weight does not include the weight of hydrazine which must be

resupplied to keep an additional 2.9 kw of solar array in' orbit.

An alternate approach would be to redesign all air handling com-ponents of the EClS to maintain required airflow while holding the fan power increase to a minimum. This solution requires larger heat exchangers. filters, and distribution air ducting, as well as larger fans. The result of a family of such system designs is shown on Figure 10-5. Note on this Figure that the fan power delta ;s now only 1.9 KW in going from a sea level to 11 psia cabin. This is preferable to the 2.9 KW delta which results from changing only the fans, as shown on Figure 10-4. This lower power penalty is obtained by increasing the size of other air handling components in the system by 28 lb. and 1.7 ft 3

10.3.3 Conclusions Regarding Selection Of Cabin Pressure

It is beyond the scope of this study to recommend the design value of cabin pressure which should ultimately be selected for SOC, but a "suggested" value is presented in this Section. As the preced-

ing sections have pointed out, there are so many diverse factor to consider that the final selection is a difficult compromise. The following is a set of individual conclusions which may be reached concerning these factors:

10-18

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a: r.J Po.

0:: r.J 5

2 :z /I( r... ...:I

~ 4

r.J Ul <

D180-26495-4

80~----~------~-----r CONSTANT CABIN

TEMPERATURE

~ 60~--------r-----l~~------~--------~------~ u :z .... a: r.J

~ Po.

:z ~

~ r.J U

40~--------+-----'---r--~~~r-------~--------~

BASELINE DESIGN POINT-

~ 20~--------r-------~------~--~~--~-------; Po.

O~---~-----r---'~--~---T--~-----r---~----r~~ 10 11 12 13 14 15

CABIN PRESSURE (PSIA)

FIGURE 10-4

CABIN PRESSURE AFFECT ON FAN POWER

10-19

~ ::.::

.ct: ~ ...:l IJJ 0

0:: IJJ ~ 0 tl.

M ~ r:...

.ct: ~ ...:l IJJ 0

IJJ ::E: ::> ...:l 0 :>

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< ~ ..:I IJJ 0

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0180-26495-4

2 ~

'" ...... i'-.. 1 " ~ ............

'-0

, , 2 ~

I" to..... 1

~ ~

'" r-..... 0 '" ~

DESIGN CABIN PRESSURE (PSIA)

FIGURE 10-5

PENALTY ON REDUCED CABIN PRESSURE (FAMILY OF OPTIMIZED SYSTEMS)

10-20

D180-26495-4

1. Referring to Figure 10-2 it can be seen that the logical cabin

pressure for SOC 1 ies within the boundaries of a triangle

formed by the 26.9 percent oxygen Fire Limit on the left, the

tentative O2 Toxicity Limit and "Practical" Suit Limit on the

right, and the 8,000 foot equivalent oxygpn level at the

bottom of the triangle. Existing Shuttle pressures are shown

on the Figure for reference.

2. TherE! is a preponderance of medical/health logic to favor

selecting the SOC cabin toward the upper right portion of the

triangular boundaries, mainly because man obviously works

best near his ancestral sea level environment. The power,

weight, and volume penalties of operating toward the upper

right portion of the triangle, as opposed to operating toward

t h el 0 w e r 1 eft po r t ion, are not g rea t The s e pen a 1 tie s are

about one percent of the total resources of SOC.

3. A normal cabin pressure error band of .:t.2 psi is recommended for SOC, based on this value being used on the current Shut-

t 1 e.

4. The boundaries of the triangle call for tighter control on

normal oxygen partial pressure level than is exercised on the

existing Shuttle. A control of about .:t.11 psia oxygen par-tial pressure is recommended for SOC. This is the band used

by the STS-l for EVA support, shown on the Fi gure, and is

held by manual control. Automatic control on SOC should be

at least this accurate.

5. The Ibox 1 abel ed "suggested for SOC" on Fi gure 10-2 is just

that, a suggested compromise between the many diverse factors

involved. Based on information ava"ilable during this study,

it is a logical, but not firm, selection. Use of the trade-

off factors presented in Section 10.3 allows evaluation of the

effect of other cabin pressure over the full range being

considered.

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10.4 REDUNDANCY PHILOSOPHY

The reliability of EClS equipment to perform its intended function is improved by installing redundant equipment. Previous manned space programs relied on this principle of installed redundancy to provide a reliability adequate to achieve their mission objec-tives. and these early space programs were of a mission length which made achieving reliability goals by this method feasible. The SOC has a 10 to 20 year expected useful life requirement. Because of this long life requirement providing adequate relia-bility by installed redundancy ;s not feasible. Hardware designed for inorbit maintenance is mandatory to achieve the long SOC mission. Maintenance, however, does not delete the requirement for needing installed redundancy, but the amount of installed redundance for SOC EelS hardware is dictated by different reasons than past manned space vehicles. The key reasons for installed redundancy on SOC are:

A fail operational/fail safe design requirement

No single EelS failure shall cause abandonment of a habitat module

No single EClS failure shall require a Shuttle flight before the next planned flight.

The fail operational/fail safe requirement dictates the need to withstand two non-maintainable worst failures and still remain in a safe operating mode. A non-maintainable worst failure does not mean to imply that the EelS system is not maintainable. All of the EClS system can be maintained. however. some of the equipment such as main distribution plumbing, major wiring distribution bundles and equipment support structure, all which have a reli-ability of nearly "one" and would be expected to last for the life

of the SOC, will be difficult to maintain and may require equip-ment and/or specialists to be supplied by the next Shuttle flight in order to conduct the maintenance. A non-maintainable failure also exists if the last spare has been used for equipment which is expected to be maintained. This first ground rule needs specified mission time periods to be meaningful. The fail operational time

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period is defined as 90 days, which is the normal SOC resupply

period. The fail safe time period is defined to be 300 hours,

which would allow for an emergency rescue by Shuttle.

The requirement that no sing'le EClS failure shall cause abandon-

ment of a habitation module makes it clear that abandoning a

habitat due to a single worst non-maintainable failure is unac-

ceptable. The requirement that no single failure shall require a

Shuttle flight before the next planned flight is self-explanatory.

As a result of the above, all critical EelS functions must: be

redundant.

The implementation of this redundancy philosophy requires the

installation of dual heat transport and rejection loops, dual air

revitalization subsystems, and dual water processing modules in

each habitat SOC.

Theine are instances, however, where dual redundancy per habitat

was not followed. For example, there are four cabin ventilation

and thermal control packages installed in each habitat module.

The sizing of this equipment into a larger number of smaller

modules "'as determined by the desire for commonal ity with the

thermal control units used in the service module. The service

module units need to handle only about one half the heat load and

ventilation flow as compared to the habitat module requirements.

Only one O2

generation and one hydrazine decomposition subsystems

were installed in each service module, which in turn supports one

habitable module. Each, however, are double sized to be capable

of servicing the full SOC. Intercabin plumbing between modules

permits 'ither of these subsystems to maintain the pressure and

atmosphere composition control in both habitat modules and both

service modules. Fail safe operation following two non-repairable

failures is provided by a 300 hour stored gaseous supply.

Some of the equipment categorized under health and hygiene are not

considered as critical functions. The backup capabil ity provided

by inflight maintenance and alternate operation modes will provide

a reliab'ility level commensurate with the mission requirements.

Therefore, one washin~ machine, one shower and one dishwasher are

considered adequate.

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A complete listing of SOC EClS system packages including the

location of the packages, and the number of packages installed in each location is provided in Table 10-2.

This redundancy philosophy which has evolved for SOC will provide a comfortable environment in each habitat for 4 man crew with short term capabi 1 ity for an 8 man crew when all of the EClS equipment is operational. After the first worst failure (before

maintenance is performed). the EClS equipment will provide an acceptable environment for a 4 man crew. Even with the fir s t

worst non-maintainable failure the EClS system will support an 8 man crew for a 300 hour emergency period at degraded levels if the other habitat must be evacuated. The capability of the system to satisfy various emergency levels is presented in Section 10.2.

10.5 MAINTENANCE CONCEPT DEFINITION

The EClS system, with its many pumps, fans, valves, instruments, controllers, etc., must be designed to be in-orbit maintainable in order to achieve the 10-20 year useful life required for the SOC.

The hardware in the EClS system will be designed for maximum life but all dynamic hardware, like that mentioned above, will have an unsatisfactory probability of failure and replacement will be required. In-orbit maintenance places specific design require-ments on the hardware. It al so must be assumed that mal ntenance must be performed by a SOC crew man who does not have the detailed training that a factory technician would have and further must be performed with the general purpose tools available in the SOC tool kit. The general requi rement for the EClS system to be in-orbit maintainable and the assumptions regarding crew training and tool availability dictate that the EelS system be designed so that:

Fault isolation to the lowest Replacement level(lRU) hardware item be generally automated (some interaction with the crew man to provide yes/no information to the fault isolation process is acceptable).

The lRU hardware item be adequately accessible.

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'l'ABLE 10-2

ECLS EQUIPMENT PACKAGES AND LOCATION

Humber of 'ackaq,o. and t.ocauon Totat SOC EeLS

tCLS rUNCTXOrl ...... JOII. tt"Utl'1'Inl'l' HAe L ""8 :2 ,. 1 ,. I '" ... PaC'acz VontlJ. .. tlon 'ana 4 4 ~ ~ I.. 112"""'.-2 ..... C.b1n VontU.tlon and Thora.l Control Air Coo1!n9 fI t Eachaf'Hjl.re 4 4 Cold Pht /0 /0 26 26 40 .. -32,"""",.

~hl,l"ldlt leatton 2 2 4 COl JlellloOval

~ ~ 4 Catalytic o.ldiur 4 An Pllvttalhation

CO 2 I'teduction 2

Odor Control 2 2 4 Ataoeptutu fmnt tortn" 1 2

~d torl 2 ~ 4 froon Cool .. nt pu.p 'edr.a,,". 2 6 ""at Tranapo1"t and l'ta)Octlon 'reon To Water .... t tach.anq'u. 2 2 4

Waur Coolant Puap '.cu, 2 2 4 0, ee"...ratlOft 2 Hydrulna o.cOIOpoett1on"":z Supply

5 2

AUIIOaphorlc: Supply "ydraun. Slot(l9_ 5 boet9.ncy 0, Stor_;_ / / taor90ney "2 Stor~. 2 2 tv.poranon p",ntlc.tion Un~t. 2 1 4 NatoI' OuaUty P')()n1torlnq 1 2

Nolt.or .. roc tnq And Mflneq __ nt .... t ..... t.r Stor.9. 3 3 g PoUblo Watar 5tor89_ :3 t:a.tqoru:y "tlt..- 8tol:ID90 /EVA Wator 4 22 26

W.ate CoUoetlon And 8tor.,_ / 2 berg"nC'y W t. Collectlon 1 2 tkJt Water Supply I ~ Cold Waul' Supply 1 Showor - 1 Hand W.ah 1 2

Ho.l th ~no Ryq i 0". ClOtho. Washor/Dryer I ~ Tr h COtIpactor 1 rood Rotrlqerot.ol' 1 2 rood Fro.aer 1 aven 2 Dlshvaahor /

5u' U And aactpacke

T T ~ JIIoet..arqo SU,tiona r:W./IVA Support Au LoC'k Support 1 1 2

t.ltrqo"c:y [licap. $yateR :3 2 5 tCI..S Co"traJ Conll'ol/tlhphy : 2 Syll". Control Portable "'alnt"nanco Contro,\/Dhplay 2

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LRU attachment fittings, plumbing connections, wire connec-

tors, etc. be generally standardized to minimize training and

tool requirements and captivated to avoid loss.

Drain and refill of systems be avoided for LRU replacement.

The definition of the LRU requires optimization studies. Pre-

vious EClS systems studies (i.e., the Space Station Prototype

program) indicated that the component level {pump, fan, valve,

instrument, controller, etc.} was the proper maintenance level.

The component level probably allows for the maximum use of common-

ality. Common components used in a variety of installations will

significantly reduce spares inventory requirements, logistics

requirements and crew training. The disadvantage is that common

components will be slightly oversized or undersized for some

installations.

With minor exception maintenance below the component level is not

practical because of excessive crew training, spares inventory and

special tool requirements. A higher level lRU than components may

have some merit. Benefits may be realized in reduced weight

abundance of installed equipment, on-board spares and resupplied

spares. Crew time and fault isolation instrumentation and soft-

ware will also be reduced. At this point a component level lRU ;s

being assumed. however. during the SOC Phase B effort when more

detailed subsystem schematics and hardware sizing information is

available, an optimization study should be conducted to determine

the appropriate lRU level and the degree to which hardware common-

ality is incorporated. Figures 10-6 and 10-7 show graphically how

an optimum lRU selection and a commonality decision might be made.

In addition to determining an optimum lRU and commonality levels,

as discussed above, certain goal s must be set in order to have a

comprehensive maintenance philosophy. The following goals, some

of which were stated earlier, have been established to minimize

the impact of I:laintenance on the SOC EClS system:

10-26

COMPONENT

0180-26495-4

SYSTEM PENALTY

-~, _______ INSTALLED

SIZE OF REPLACEMENT UNIT

FIGURE 10-6

EQUIPMENT

ON-BOARD SPARES

RESUPLY SPARES

SUBSYSTEM

LOWEST REPLACEABLE UNIT OPTIMIZATION

10-27

.. ~ l? H W ~

0180-26495-4

DIVERSITY OF SPARES

-C COMMONALITY

FIGURE 10-7

LOWEST REPLACEABLE UNIT COMMONALITY OPTIMIZATION

10-28

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No maintenance task shall exceed 4 hours.

Unscheduled maintenance shall not exceed 8 man hours per month.

Scheduled maintenance shall not exceed 40 man hours per month.

Maintenance shall be accomplished with SOC tool kit TBD.

Maintenance skill level shall be within SOC technician cap-ability.

Spares commonality shall be selected to provide minimum penalty.

Equipment shall be designed to avoid fluid loss and air inclusion during maintenance.

For ease of maintenance, life support equipment will be located behind removable ceiling and floor panels. The equipment will be located so that repl aceabl e component and/or modul ar units are generally no more than a single layer deep with adequate perimeter access py'oviding up to five side access for maintenance. The replaceable modular units are packaged within a panel cavity so that plenum airflow is not disturbed during a maintenance pro-cedure.

Modular units should generally be configured as a group of com-ponents forming a logic group. On-board repair of components is not ordinarily considered feasible. However, consideration should be given to possible emergency repairs below the component level by using standard (common) parts.

Malfunctions shall be isolatable to the logic group level and, as a 90al, to the LRU level. Automatic detection of the' malfunction or degradation shall be provided for critical functions. Where the fault isolation cannot be narrowed to a specific LRU in the allowable down time, replacement shall be made to all suspect LRU's.

10-29

Sufficient spares shall performance within the ments. Cannibalization is possible to be done

0180-26495-4

be stored on-board to sustain the system resupply period, and reliability allot-of other systems can be considered if it within allowable maintenance time period.

The components in the SOC system which require replacement fall into several combinations of the following categories: low pres-

sure, high pressure, hazardous liquid, non-hazardous liquid, hazardous gas, and non-hazardous gas.

Gas-line components. In general, gas-line component problems are not as severe as liquid-line problems. Some gas lines do, how-ever. contain hazardous gases (such as hydrogen) or contaminated gases (such as commode outlet gas). Lines carrying hazardous gases have provisions for either purging or evacuating prior to component removal. Thereafter, maintenance considerations are the same as for any other gas-line and are described in the following paragraphs.

High-pressure gases. High-pressure lines are defined generally as those containing pressures which exceed 5 psig, while low-pressure

lines are defined generally as those containing pressures below 5 psig. Vacuum lines should be considered and maintained in the same manner as high-pressure lines.

The high-pressure gas-lines shall contain bypass lines and shutoff and depressurization valves which are properly located to allow for depressurization and maintenance without interrupting other critical system functions. Once depressurized, the 1 ines may be opened at the component fittings to allow component replacement.

Low-pressure gases. Components in the low-pressure lines will be connected to the duct through the use of flexible hoses, beaded tubes, and flanges or Marman-type couplings (or both). The Marman type flanges shall use of dovetailed grooves to captivate any seals which are used, and they should be coupled together with a quick release clamp. Where the system must remain operating, caps shall be provided to close off the open ports.

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~id-line components. When considering the liquid subsystems and the requirement that the EelS must operate in a zero-g envi-

ronment, two important aspects of liquid-line component mainten-

anCE! became apparent: (a) prevention of liquid loss and, (b)

prevention of gas ingestion. Prevention of 1 iquid loss precludes

a recharge operation with the attendant need for replacement

1 iquid and minimizes spillage which could induce other component

failures, introduce possible contamination, and require varied,

complex, a~d time consuming cleanup operations. Prevention of gas ingestion eliminates the need for evacuation or bleeding and

reduces the performance requi rement of gas separat ion equi prnent.

One of the maintenance methods considered for liquid-line compon-

ents was the subsystem "drain and fill" approach. Although this

approach would permit the use of a greater number of "off-the-

shelf" or standard components, the design constraint resulting from

the possibility of gas entrapment could jeopardize the system

integration effort. This approach would also require many drain

ports and complex servicing equipment.

Requirements for draining, complex servicing equipment, and clean-

up problems are undesirable. Therefore, the subsystem shutdown

and drainage approach has been eliminated as a normal liquid-line

maintenance approach. The selected basic approach to maintenance

of the EelS liquid-line subsystems is that it can be accomplished

without requlrlng draining or complete subsystem deactivation.

Sevleral concepts are being developed which allow for replacement

of components and valves without system drainage and also allow

for bypassing failed line sections.

Non-maintainable Items

Failure of an item which is not normally maintainable is consid-

ered an exceptional event. Repair of these items may require disruption of normal activities since the repair effort may be of

long duration, requiring more than one crew member, and, in ex-

treme, the evacuation of thE! module may be required. The first

worst case failure of this equipment will cause a degraded perfor-

mance condition to exist which is acceptable for 90 day operation.

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Items which cannot be maintained by the use of on-board spare

parts are requi red to have a rel iabil ity which approaches one. This equipment includes frames, main distribution fluid lines and

main distribution electrical harnesses.

Non-maintainable item should be protected so that maintenance of adjacent components does not pose the possibility of damage to the non-maintainable item. In the case of fluid lines and harnesses, these should be located behind panels and the harnesses should be routed in conduit. Where redundant fluid lines and harnesses are required, they should be routed in separate compartments and

should be located as far apart from one another as possible.

Non-maintainable items should be designed with conservative safety factors. possible.

Major structural members should contain redundancy where The non-maintainable items should be designed so that a

failure can be easily detected.

In order to minimize the number of random failures, low reliabil-ity and life limited items should be replaced before the expected failure occurs. The tradeoff is between crew convenience and

cost. The minimum cost approach is to let all equipment operate until it wears out. The maximum cost approach is to replace equipment at the minimum expected life. A third approach is to monitor equipment and predict equipment failure based on trend

analysis. If the predictions are fairly accurate, this third method would be close to the minimum cost, while providing flex-ibility for scheduling maintenance. The third approach should be considered whenever it can be accomplished without the addition of instrumentation. It should also be considered when the equipment is costly and there is a large spread between the minimum and maximum expected life.

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10.6 EVALUATION OF SELECTABLE CABIN TEMPERATURE, AND CABIN TEMPERATURE BAND EXPANSION

There are several facets to a discussion of variations in cabin temperatul~e. First, there is the matter of control band varia-tion, that is to say the cabin temperature band over which the cabin temperature controls may be set by the crew. The set tem-perature is not related to the capacity of the EClS system to always provide this selected temperature. For example, the ther-mostat us.ed by the crew to select cabin temperature could be

settlable at any temperature between 65F and 80F. When in the full cold or full hot mode, any cabin temperature within this band would be maintained within the capacity of the EClS to do so. Cabin cooling capability, as discussed in this section, refers only to the full cold capability of the EClS to satisfy the oper-ating condition in question. The degree to which cooling capacity should be provided, to ensure that the lowest cabin temperature selectabll:? by the crew will be achieved, has weight, volume. and power impact on the EClS system. On the other hand, the ability to heat the cabin has no penalty since the vehicle is so well insulated, and since there is always a significant electrical

heat load to be dissipated. It is the cooling capacity required that sizes the thermal control packages, and the remainder of the heat rejection system as well.

This SOC study considers as its baseline a system capable of providing the following "full cold" cabin cooling capacity:

No. Temp. Hab. Mod. Control Crew Per Maximum Units Operating "Full Maximum

Cabin EClS Fa il ed Temp. Cold" Cabin Heat Crew Per Operating Per' Control Cabin Dew Load Hab. Mod. Mode Hab. Mod. Unit Temp. 'P'oi nt

Maximum 4 Operational 0 1 75 60

Maximum 4 90 Day 2 2 85 70 Degraded

Maximum 8 90 Day a 2 85 70 Degraded

Maximum 8 14 Day 2 4 90 75 Emergency

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In practice, the thermal system cannot be sized to simultaneously "just meet" all four of the above operating modes. In this case, the operational mode will be just met, and the cooling capacity requirement of the other three will be somewhat exceeded. The baseline system of this study, in the operational mode, will provide sufficient full cold cooling capacity at the maximum heat load case for a cabin temperature of 75F. The question arises;

what would be the effect on the EClS system if this cool ing ca-pacity requirement were increased to drop this maximum operational cabin temperature to 70F, and 65F respectively.

Preliminary designs of a family of different systems would exhibit the power, weight and volume penalties shown o~ Figure 10-8. Note that the penalties in gOing from 70F to 65F are much more severe than going from 75F to 70F. The reason for this is that there is an i nfi ni te penal ty in goi ng to a 55F cab; n because the ability to hold cabin relative humidity of 75% becomes impossible below a cabin temperature of 550'F. This is because the coolant water and therefore the dew point is 1 imited to 40F to prevent icing in the coolant water to freon heat exchanger.

The penalty study results shown in Figure 10-8 indicate that there is relatively little penalty to change the present 75F selected SOC baseline cabin temperature specification to require a full cold operational maximum temperature of 70F. This figure sho\,/s that the increased comfort which would result from a 70F cabin at high work rates would cost approximately 35 lb, 2.2 ft 3 and 400 watts. A further decrease in full cold cabin temperature capa-bility to 65F appears to have an unacceptable penalty.

This study does not recommend a reduction in max load cabin tem-perature capability from the baseline 75F value to .~OF, but the tradeoff study suggests that the increased penalty of this 5F

reduction is small enough to make such a change in the specifica-tion reasonable if it is desired for additional comfort for high work loads.

10-34

~ 2.0

1.5 ~ E-t 1.0 ...:I W Cl

0.5 Il:: w

0 ~ Il. -0.5

M 16 E-t Ii.

12

8

4

o -4

: 400 III ...:I

~ E-t ...:I W Cl

E-t :x: t? H

300

200

100

o

r\

"

1\ \

\ \

~ -100 65

"-

t\.

"

'" "-

D180-26495-4

'" r--..... ......... t-- t-- "- r-- ....

~ ..... ...... ... --~ -, NO'l'E: RADIATOR

NOT INCLUDE D

"-BASELINE DESIGN POINT-I

./

.......... r ........ ,/ V "-

70 75 DESIGN MAXIMUM OPERATIONAL

CABIN TEMPERATURE (OF)

FIGURE 10-8

80

SELECTABLE CABIN TEMPERATURE PENALTY

10-35

D180-26495-4

Most buildirig air conditioning systems include a thermostat setta-ble in the range of 65F to 80F, regardless of the capacity of the buildings system to satisfy this range under conditions of maximum load. It is suggested that the SOC cabin temperature

thermostat likewise have a thermostat settable range from 65F to

80F.

10.7 EClS SUBSYSTEM SELECTION RATIONALE

At the initiation of Boeing1s SOC Systems Analysis effort, an EClS reference baseline was defined by NASA in NASA-3 and NASA-6. The

EClS system defined in Boeing-19 differs from the NASA reference

basel ine in some areas. A rationale for the most significant

deviations is provided in this section. The differences discussed are:

CO 2 Removal Concept Selection Wastewater Processing Concept Selection

10.7.1 CO 2 Removal Subsystem Concept Selection

Hamilton Standard has been evaluating, by spacecraft CO 2 removal, concepts for more than

pendable and regenerable techniques have been cepts receiving considerable attention were:

lithium hydroxide Lithium peroxide Solid amines Molecular sieves Electrochemical Molten carbonate

analysis and test, 20 years. Both ex-

evaluated. The con-

With the exception of Skylab, all U.S. spacecraft to date have

used expendable lithium hydroxide. Skylab used a regenerable molecular sieve CO 2 removal and dump system. Spacecraft such as the pl anned SOC wi 11 requi re a regenerabl e CO 2 removal system which will supply pure CO 2 at atmospheric pressure for a CO 2 re-duction process in order to save the oxygen. Two concepts have evolved which meet these SOC requirements: solid amine and elec-

trochemical. Both of these concepts have been thoroughly tested

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0180-26495-4

as prototype hardware under both NASA contracts and contractor R&D funds demonstrating their capability to perform the CO 2 removal and concentration functions. For the SOC study activity the Solid Amine Water (Atmospheric Pressure Steam) Desorbed (SAWD) has been selected by Hamilton Standard. The reasons for this selection are that SAWD as compared to the Electrochemical Depolarized CO 2 Concentrator (EOC) when integrated into the SOC life support system:

1. Has less system weight, volume and power penalty.

2. Is not dependent on an emergency liOH backup system.

3. Does not require over-sizing the electrolysis subsystem.

4. Is less dependent on other subsystems for operation (cas-cading failures).

5. Allows the EClS to be designed without introducing H2 into habitat module or requiring H2 lines passing through bulk-heads.

6. Is free from potential caustic carryover.

7. Is tolerant to the cabin humidity range without precondi-tioning the air.

8. Can be exposed to a depressurized cabin without requiring shut-off valves or other precautionary action.

9. Uses power on the light side of the orbit without a signi-ficant sizing penalty.

10. Is less expensive hardware.

In order to evaluate statement number 1, it is necessary to pro-vide a dE~sign specification for sizing the SOC CO 2 removal unit. It is believed that as a result of the SOC fail operational (90 day degraded)/fail safe (14 day further degraded) requirement the SAWD

CO 2 removal units must be s'ized for 8 men CO 2 output at 12 mmHg partial pressure. Four un'its sized to this criteria would be installed in the SOC, two per habitat, to meet safety require-ments. The EDC CO 2 removal unit can be sized for the same case as

10-37

0180-26495-4 the SAWD supported by four 4 man (+ EDC 02 and H2 ) electrolysis units operated on the sun side only. In this case. EDC would shut down on the dark side of the orbit. As another option, the EDC can be sized to run continuously with continuously running elec-trolysis units. In order to meet safety requirements. two EDC units and two electrolysis units would be required per habitabil-ity module in both of these cases.

Another approach using an EDC unit which would also meet the safety criteria is to incorporate a backup LiOH s